EP2788609A1 - Moteur à turbine à gaz possédant une buse à surface variable de ventilateur mobile axialement - Google Patents

Moteur à turbine à gaz possédant une buse à surface variable de ventilateur mobile axialement

Info

Publication number
EP2788609A1
EP2788609A1 EP12869088.0A EP12869088A EP2788609A1 EP 2788609 A1 EP2788609 A1 EP 2788609A1 EP 12869088 A EP12869088 A EP 12869088A EP 2788609 A1 EP2788609 A1 EP 2788609A1
Authority
EP
European Patent Office
Prior art keywords
fan
nacelle
assembly
engine
recited
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP12869088.0A
Other languages
German (de)
English (en)
Other versions
EP2788609A4 (fr
Inventor
Gregory A. Kohlenberg
Sean P. Zamora
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Raytheon Technologies Corp
Original Assignee
United Technologies Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Priority claimed from US13/314,365 external-priority patent/US9701415B2/en
Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP2788609A1 publication Critical patent/EP2788609A1/fr
Publication of EP2788609A4 publication Critical patent/EP2788609A4/fr
Withdrawn legal-status Critical Current

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/075—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type controlling flow ratio between flows
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/06—Varying effective area of jet pipe or nozzle
    • F02K1/09—Varying effective area of jet pipe or nozzle by axially moving an external member, e.g. a shroud
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan

Definitions

  • the present invention relates to a gas turbine engine, and more particularly to a turbofan engine having a fan variable area nozzle (VAFN) which moves axially to change a bypass flow path area thereof.
  • VAFN fan variable area nozzle
  • Conventional gas turbine engines generally include a fan section and a core engine with the fan section having a larger diameter than that of the core engine.
  • the fan section and the core engine are disposed about a longitudinal axis and are enclosed within an engine nacelle assembly.
  • Combustion gases are discharged from the core engine through a core exhaust nozzle while an annular fan flow, disposed radially outward of the primary airflow path, is discharged through an annular fan exhaust nozzle defined between a fan nacelle and a core nacelle.
  • a majority of thrust is produced by the pressurized fan air discharged through the fan exhaust nozzle, the remaining thrust being provided from the combustion gases discharged through the core exhaust nozzle.
  • the fan nozzles of conventional gas turbine engines have a fixed geometry.
  • the fixed geometry fan nozzles are a compromise suitable for take-off and landing conditions as well as for cruise conditions.
  • Some gas turbine engines have implemented fan variable area nozzles.
  • the fan variable area nozzle provide a smaller fan exit nozzle diameter during cruise conditions and a larger fan exit nozzle diameter during take-off and landing conditions.
  • Existing fan variable area nozzles typically utilize relatively complex mechanisms that increase overall engine weight to the extent that the increased fuel efficiency therefrom may be negated.
  • a nacelle assembly for a high -bypass gas turbine engine may include a core nacelle defined about an engine centerline axis, a fan nacelle mounted at least partially around the core nacelle to define a fan bypass flow path for a fan bypass airflow, and a fan variable area nozzle axially movable relative the fan nacelle to define an auxiliary port to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation.
  • the controller may be operable to control the fan variable area nozzle to vary a fan nozzle exit area and adjust the pressure ratio of the fan bypass airflow.
  • the controller may be operable to reduce the fan nozzle exit area at a cruise flight condition.
  • the controller may be operable to control the aid fan nozzle exit area to reduce a fan instability.
  • the fan variable area nozzle may define a trailing edge of the fan nacelle.
  • the assembly may further include a controller operable to axially move the fan variable area nozzle to vary the fan nozzle exit area in response to a flight condition.
  • the fan variable area nozzle may be aligned with the fan nacelle to define a closed position of the fan nozzle exit area. Additionally or alternatively, the fan variable area nozzle is axially offset from the fan nacelle to define an open position of the fan nozzle exit area.
  • the nacelle assembly may further include a gear system driven by the core engine within the core nacelle to drive the fan within the fan nacelle, the gear system defines a gear reduction ratio of greater than or equal to about 2.3.
  • the nacelle assembly may further include a gear system driven by the core engine within the core nacelle to drive the fan within the fan nacelle, the gear system defines a gear reduction ratio of greater than or equal to about 2.5.
  • the nacelle assembly may further include a gear system driven by the core engine to drive the fan, the gear system defines a gear reduction ratio of greater than or equal to 2.5.
  • the core engine may include a low pressure turbine which defines a pressure ratio that is greater than about five (5).
  • the core engine may include a low pressure turbine which defines a pressure ratio that is greater than five (5).
  • bypass flow may define a bypass ratio greater than about six (6).
  • the bypass flow may define a bypass ratio greater than about ten (10).
  • bypass flow may define a bypass ratio greater than ten (10).
  • a gas turbine engine may include a core nacelle defined about an engine centerline axis, a fan nacelle mounted at least partially around the core nacelle to define a fan bypass flow path for a fan bypass airflow; a fan variable area nozzle axially movable relative the fan nacelle to define an auxiliary port to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation, and a controller operable to control the fan variable area nozzle to vary a fan nozzle exit area and adjust the pressure ratio of the fan bypass airflow.
  • the gas turbine engine may be a direct drive turbofan engine.
  • the gas turbine may further include a low spool within the core nacelle that drives a fan within the fan nacelle through a geared architecture.
  • the engine may have a bypass ratio greater than 10: 1 and the geared architecture may have a gear reduction ratio of greater than 2.5:1.
  • Figure 1A is a general schematic partial fragmentary view of an exemplary gas turbine engine embodiment for use with the present invention
  • Figure IB is a rear view of the engine
  • Figure 1C is a side view of the engine integrated with a pylon
  • Figure ID is a perspective view of the engine integrated with a pylon
  • Figure 2A is a sectional side view of the VAFN in a closed position
  • Figure 2B is a sectional side view of the VAFN in an open position
  • Figure 3 is a graph of a bypass duct normalized cross-sectional area distribution.
  • Figure 4 is a graph of a Effective Area Increase vs. Nozzle Translation
  • Figure 5 is a graph of a duct area distribution
  • Figure 6A is schematic geometric view of the auxiliary port location
  • Figure 6B is schematic geometric view of the auxiliary port entrance angle
  • Figure 6C is schematic geometric view of a VAFN outer surface curvature.
  • Figure 1A illustrates a general partial fragmentary schematic view of a gas turbofan engine 10 suspended from an engine pylon P within an engine nacelle assembly N as is typical of an aircraft designed for subsonic operation.
  • the turbofan engine 10 includes a core engine within a core nacelle 12 that houses a low spool 14 and high spool 24.
  • the low spool 14 includes a low pressure compressor 16 and low pressure turbine 18.
  • the low spool 14 drives a fan section 20 through a gear train 22.
  • the high spool 24 includes a high pressure compressor 26 and high pressure turbine 28.
  • a combustor 30 is arranged between the high pressure compressor 26 and high pressure turbine 28.
  • the low and high spools 14, 24 rotate about an engine axis of rotation A.
  • the engine 10 is preferably a high-bypass geared architecture aircraft engine.
  • the engine 10 bypass ratio is greater than about six (6) to ten (10)
  • the gear train 22 is an epicyclic gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 18 has a pressure ratio that is greater than about 5.
  • the engine 10 bypass ratio is greater than ten (10: 1)
  • the turbofan diameter is significantly larger than that of the low pressure compressor 16, and the low pressure turbine 18 has a pressure ratio that is greater than 5: 1.
  • the gear train 22 may be an epicycle gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than 2.5: 1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
  • the fan section 20 communicates airflow into the core nacelle 12 to power the low pressure compressor 16 and the high pressure compressor 26.
  • Core airflow compressed by the low pressure compressor 16 and the high pressure compressor 26 is mixed with the fuel in the combustor 30 and expanded over the high pressure turbine 28 and low pressure turbine 18.
  • the turbines 28, 18 are coupled for rotation with, respective, spools 24, 14 to rotationally drive the compressors 26, 16 and through the gear train 22, the fan section 20 in response to the expansion.
  • a core engine exhaust E exits the core nacelle 12 through a core nozzle 43 defined between the core nacelle 12 and a tail cone 32.
  • the core nacelle 12 is supported within the fan nacelle 34 by structure 36 often generically referred to as Fan Exit Guide Vanes (FEGVs).
  • a bypass flow path 40 is defined between the core nacelle 12 and the fan nacelle 34.
  • the engine 10 generates a high bypass flow arrangement with a bypass ratio in which approximately 80 percent of the airflow entering the fan nacelle 34 becomes bypass flow B.
  • the bypass flow B communicates through the generally annular fan bypass flow path 40 and is discharged from the engine 10 through a fan variable area nozzle (VAFN) 42 which defines a fan nozzle exit area 44 between the fan nacelle 34 and the core nacelle 12 at a fan nacelle end segment 34S of the fan nacelle 34 downstream of the fan section 20.
  • VAFN fan variable area nozzle
  • Thrust is a function of density, velocity, and area. One or more of these parameters can be manipulated to vary the amount and direction of thrust provided by the bypass flow B.
  • the VAFN 42 operates to effectively vary the area of the fan nozzle exit area 44 to selectively adjust the pressure ratio of the bypass flow B in response to a controller C.
  • Low pressure ratio turbofans are desirable for their high propulsive efficiency. However, low pressure ratio fans may be inherently susceptible to fan stability/flutter problems at low power and low flight speeds.
  • the VAFN allows the engine to change to a more favorable fan operating line at low power, avoiding the instability region, and still provide the relatively smaller nozzle area necessary to obtain a high-efficiency fan operating line at cruise.
  • the fan section 20 of the engine 10 is preferably designed for a particular flight condition— typically cruise at about 0.8M and about 35,000 feet.
  • the VAFN 42 is operated to effectively vary the fan nozzle exit area 44 to adjust fan bypass air flow such that the angle of attack or incidence on the fan blades is maintained close to the design incidence for efficient engine operation at other flight conditions, such as landing and takeoff to thus provide optimized engine operation over a range of flight conditions with respect to performance and other operational parameters such as noise levels.
  • the VAFN 42 is separated into at least two sectors 42A-42B ( Figure IB) defined between the pylon P and a lower Bi-Fi splitter L which typically interconnects a larger diameter fan duct reverser cowl and a smaller diameter core cowl ( Figure 1C and ID).
  • Each of the at least two sectors 42A-42B are independently adjustable to asymmetrically vary the fan nozzle exit area 44 to generate vectored thrust. It should be understood that although two segments are illustrated, any number of segments may alternatively or additionally be provided.
  • the VAFN 42 communicates with a controller C or the like to adjust the fan nozzle exit area 44 in a symmetrical and asymmetrical manner.
  • Other control systems including an engine controller or aircraft flight control system may also be usable with the present invention.
  • thrust efficiency and fuel economy are maximized during each flight condition.
  • circumferential sectors 42A-42B of the VAFN 42 to provide an asymmetrical fan nozzle exit area 44, engine bypass flow is selectively vectored to provide, for example only, trim balance or thrust controlled maneuvering enhanced ground operations or short field performance.
  • the VAFN 42 generally includes an auxiliary port assembly 50 having a first fan nacelle section 52 and a second fan nacelle section 54 movably mounted relative the first fan nacelle section 52.
  • the second fan nacelle section 54 axially slides along the engine axis A relative the fixed first fan nacelle section 52 to change the effective area of the fan nozzle exit area 44.
  • the second fan nacelle section 54 slides aftward upon a track fairing 56A, 56B (illustrated schematically in Figure 1C and ID) in response to an actuator 58 (illustrated schematically).
  • the track fairing 56A, 56B extend from the first fan nacelle section 52 adjacent the respective pylon P and the lower Bi-Fi splitter L ( Figure ID).
  • the VAFN 42 changes the physical area and geometry of the bypass flow path 40 during particular flight conditions.
  • the bypass flow B is effectively altered by sliding of the second fan nacelle section 54 relative the first fan nacelle section 52 between a closed position ( Figures 2A) and an open position ( Figures 2B).
  • the auxiliary port assembly 50 is closed by positioning the second fan nacelle section 54 in-line with the first fan nacelle section 52 to define the fan nozzle exit area 44 as exit area F0 ( Figure 2A).
  • the VAFN 42 is opened by moving the second fan nacelle section 54 aftward along the track fairing 56A, 56B away from the first fan nacelle section 52 to open an auxiliary port 60 which extends between the open second fan nacelle section 54 relative the first fan nacelle section 52 to essentially provide an increased fan nozzle exit area 44 exit area Fl. That is, the exit area Fl with the port 60 is greater than exit area F0 ( Figure 2B).
  • the auxiliary port 60 is incorporated into the exhaust system of a high bypass ratio commercial turbofan engine within the bypass duct aft of the Fan Exit Guide Vanes (FEGVs; Figures 2A, 2B).
  • the auxiliary port 60 is located in the aft section of the bypass duct outer wall.
  • the bypass duct area distribution, the effective area increase vs. translation ( Figure 4), area distribution ( Figure 5), and auxiliary port 60 location ( Figure 6A) and wall curvatures (Figure 6B-6C) are tailored to provide a proper flow-field that allows the auxiliary port 60 to obtain the required additional effective exit area.
  • the auxiliary port 60 will essentially double the effective area gain due to translation.
  • the auxiliary port 60 provides a relatively low weight method of providing increased exit area to control the fan operating line without causing high system losses or unacceptable aircraft installation issues.
  • the auxiliary port exit plane 44B (defined as the plane between the stationary section's trailing edge and the moving sections leading edge) initially has an opening in which the exit plane normal vector is near-axial, but as the stroke increases, the normal vector becomes more inclined and approaches a near-radial vector. Once the exit plane normal has become near- radial, the maximum auxiliary port effectiveness has been reached. Once this point is reached, the rate of the effective area vs. translation changes from steep slope of the "well designed port” the shallow rate of the "main nozzle only", since additional area will be provided through the main nozzle 44A due to the inward slope of the core nacelle 12. A well designed auxiliary port nozzle will achieve approximately +25% effective area before the port effectiveness limit is reached.
  • the auxiliary port doubles the rate of additional effectiveness. Outside of this range, the rate of additional effectiveness may be equivalent to a translating nozzle that has no auxiliary port. Or put another way, the auxiliary port reduces the stroke necessary for a pure translating nozzle to achieve a desired effective area.
  • the cross- sectional area at the auxiliary port 60 is greater than the maximum required effective area of the VAFN 42 and the bypass duct area distribution is tailored to ensure the duct cross-sectional area forward of the auxiliary port 60 is greater than the port opening cross-sectional area. This avoids a situation where an upstream internal cross-section becomes the controlling flow area (i.e. is smaller than the exit area), which can lead to operational limits and structural issues.
  • the auxiliary port 60 in the disclosed embodiment is located no more forward than 0.1 DEL_X / L_DUCT defined from a point D at the largest radius Rmax of the annular fan bypass flow path 40 defined by the second fan nacelle section 54.
  • Rmax is defined through point D and perpendicular to the engine axis A.
  • Point D in the disclosed non limiting embodiment is located on an inner wall surface 541 of the second fan nacelle section 54 when the second fan nacelle section 54 is in a closed position.
  • DEL_X is the axial distance to the forward most point of the auxiliary port 60 from Rmax.
  • L_DUCT is the overall axial length of the annular fan bypass flow path 40. .
  • the angle between the mean port line and the fan duct outer wall is relatively low to provide well-behaved, low loss exit flow.
  • the auxiliary port 60 entrance angle (Theta_in) relative to the fan bypass duct OD wall is less than 20 degrees ( Figure 6B) while the outer VAFN surface has an R_ARC / CHORD > 0.7 where R_ARC is a radial distance from the engine axis A to a radial outer wall surface 540 of the second fan nacelle section 54 and CHORD is the chord length of the second fan nacelle section 54. ( Figure 6C).
  • the curvature of the outer wall surface 540 near the auxiliary port 60 promotes flow through the auxiliary port 60.
  • the stroke of the second fan nacelle section 54 necessary to obtain an additional 20% effective exit area is approximately 8.4 inches.
  • the VAFN 42 communicates with the controller C to move the second fan nacelle section 54 relative the first fan nacelle section 52 of the auxiliary port assembly 50 to effectively vary the area defined by the fan nozzle exit area 44.
  • Various control systems including an engine controller or an aircraft flight control system may also be usable with the present invention.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Control Of Turbines (AREA)

Abstract

Un ensemble nacelle pour un moteur à turbine à gaz à grande dérivation comprend une buse à surface variable de ventilateur mobile axialement par rapport à la nacelle du ventilateur pour définir un orifice auxiliaire afin de faire varier la surface de sortie de la buse du ventilateur et d'ajuster un rapport de pression du flux d'air de dérivation du ventilateur pendant le fonctionnement du moteur.
EP12869088.0A 2011-12-08 2012-12-07 Moteur à turbine à gaz possédant une buse à surface variable de ventilateur mobile axialement Withdrawn EP2788609A4 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US13/314,365 US9701415B2 (en) 2007-08-23 2011-12-08 Gas turbine engine with axial movable fan variable area nozzle
PCT/US2012/068336 WO2013126123A1 (fr) 2011-12-08 2012-12-07 Moteur à turbine à gaz possédant une buse à surface variable de ventilateur mobile axialement

Publications (2)

Publication Number Publication Date
EP2788609A1 true EP2788609A1 (fr) 2014-10-15
EP2788609A4 EP2788609A4 (fr) 2015-08-26

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP12869088.0A Withdrawn EP2788609A4 (fr) 2011-12-08 2012-12-07 Moteur à turbine à gaz possédant une buse à surface variable de ventilateur mobile axialement

Country Status (4)

Country Link
EP (1) EP2788609A4 (fr)
CN (1) CN104040158A (fr)
SG (1) SG11201402854VA (fr)
WO (1) WO2013126123A1 (fr)

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN108005812B (zh) * 2017-12-04 2019-06-18 中国航空发动机研究院 采用自适应机匣和自适应风扇的智能发动机
US11306681B2 (en) * 2019-01-15 2022-04-19 The Boeing Company Sheared exhaust nozzle
CN112455699B (zh) * 2020-11-13 2024-01-02 中国航空工业集团公司沈阳飞机设计研究所 一种高融合飞机后体

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US5655360A (en) * 1995-05-31 1997-08-12 General Electric Company Thrust reverser with variable nozzle
US6732502B2 (en) * 2002-03-01 2004-05-11 General Electric Company Counter rotating aircraft gas turbine engine with high overall pressure ratio compressor
US6901739B2 (en) * 2003-10-07 2005-06-07 General Electric Company Gas turbine engine with variable pressure ratio fan system
US7845902B2 (en) * 2005-02-15 2010-12-07 Massachusetts Institute Of Technology Jet engine inlet-fan system and design method
WO2008045058A1 (fr) * 2006-10-12 2008-04-17 United Technologies Corporation Gestion de la ligne de fonctionnement d'un compresseur basse pression dans un moteur de turboréacteur à double flux
WO2008045049A1 (fr) * 2006-10-12 2008-04-17 United Technologies Corporation Turbine à gaz à tuyère de section variable par déplacement axiale de la soufflante
FR2910937B1 (fr) * 2007-01-02 2009-04-03 Airbus France Sas Nacelle de reacteur d'aeronef et aeronef comportant une telle nacelle
US7950237B2 (en) * 2007-06-25 2011-05-31 United Technologies Corporation Managing spool bearing load using variable area flow nozzle
US8074440B2 (en) * 2007-08-23 2011-12-13 United Technologies Corporation Gas turbine engine with axial movable fan variable area nozzle
US8973364B2 (en) * 2008-06-26 2015-03-10 United Technologies Corporation Gas turbine engine with noise attenuating variable area fan nozzle
US8739515B2 (en) * 2009-11-24 2014-06-03 United Technologies Corporation Variable area fan nozzle cowl airfoil

Also Published As

Publication number Publication date
CN104040158A (zh) 2014-09-10
EP2788609A4 (fr) 2015-08-26
WO2013126123A1 (fr) 2013-08-29
SG11201402854VA (en) 2014-07-30

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